The invention generally relates to an aircraft gas turbine engine apparatus for bleeding boundary layer air from the wings and nacelles in order to reduce drag. In particular, the invention relates to an engine driven electrically powered apparatus for bleeding the boundary layer air and using the boundary layer bleed air for aircraft's environmental control system (ECS).
Aircraft aerodynamic drag poses a fuel consumption problem for aircraft designers. Aerodynamic drag causes a significant increase in an aircraft's specific fuel consumption and one component of the aerodynamic drag is boundary layer drag which is associated with engine nacelles, wing, pylons and other surfaces exposed to the free stream velocity which is about 600 miles per hour at cruise. As air flows on to and over a surface such as an engine nacelle it progressively builds up a low velocity boundary layer of increasing thickness. Within this boundary layer a portion of the velocity component of free stream total pressure is converted to increased static pressure. As the result of rise in static pressure, boundary layer thickness, and diffusion a point is reached where back pressure causes an otherwise laminar boundary layer to become turbulent.
Within the turbulent region a considerable amount of total pressure is converted to static temperature represented thermodynamically as an increase in entropy. By the time the boundary layer leaves the surface, or in the particular case of an aircraft gas turbine engine the end of the nacelle, an unrecoverable loss in total pressure has occurred The large entropy rise associated with turbulence is at the expense of air momentum. Turbulence also gives rise to increased static pressure which may increase the intensity of rearward acting pressure force on the surface. Now if the boundary layer thickness is kept small, separation and turbulence will not occur and drag can be substantially reduced.
One way to avoid increase in boundary thickness is to pump or bleed off boundary layer air through holes in the surface. Boundary layer pumps or compressors would be desirable from an aerodynamic standpoint but causes design problems related to weight and complexity because of the relatively large air flow rates associated with effective boundary layer pumping or bleeding. The concept has not been used for modern aircraft and engines because heretofore the specific fuel consumption benefits have been outweighed by the penalties associated with the added weight and complexity of proposed systems. One problem, addressed by the present invention, is that a boundary layer bleed system requires a great deal of extra equipment, particularly compressors, for drawing off or bleeding the boundary layer. Therefore the present invention proposes a solution to effectively reduce aerodynamic drag using boundary layer bleed with a minimal increase in the weight and complexity of the aircraft.
Modern day aircraft use gas turbine engines, which in addition to propulsion, provide secondary functions required by the aircraft systems. These secondary functions include electrical power, hydraulic power and aircraft bleed air. Bleed air is normally taken from the engine compressor and, after pre-cooling with engine fan air in a heat exchanger, is delivered to various aircraft systems such as the cowl and wing anti-ice system and the cabin pressurization and environmental control system for controlling cabin air freshness and temperature. These two systems are generally referred to as the anti-ice and ECS systems respectively. As part of the aircraft air ducting system, air is also routed in reverse flow to the engine where it powers an air turbine engine starter. Air for engine starting can be obtained from a ground cart, an on-board auxiliary power unit or bleed air from another engine.
Extraction of aircraft bleed air from the engine compressor has adverse affects on the propulsion cycle and engine life. Air taken into the engine compressor incurs a ram drag penalty (loss of momentum). Engine net thrust is equal to engine exhaust momentum minus inlet ram drag. Engine turbine power is needed to compress air and account for compressor inefficiency. Therefore, extra fuel consumption is always associated with bleed air (air which does not produce thrust).
This extra fuel burned in the engine combustor results in higher gas temperature delivered to the engine turbine and reduction of turbine blade life. Such penalties must be incurred in order that the engine turbine provide extra power associated with bleed air. It is not possible, without undue complexity, to always bleed the engine compressor stage which provides exactly the correct pressure needed for the aircraft anti-ice and ECS systems. Typically only two bleed ports are provided. Therefore, the result is to bleed air which exceeds minimum pressure requirements resulting in even higher penalty to the engine cycle than would be required by the aircraft systems. Most often the bleed air is not only at a higher than required pressure, it is also too hot. For reasons of fire safety, maximum bleed air temperature is usually limited to 450.degree. to 500.degree. F.
Temperature control requires cooling the bleed air with a pre-cooler. Most modern engines use fan air to cool compressor bleed air. Use of fan air imposes an additional penalty on fuel consumption. Further, the pre-cooler is usually large and requires a fan air scoop which produces drag. A typical large turbofan engine will consume about 3% extra fuel and run at about 50.degree. F. hotter turbine temperature in order to provide aircraft system bleed air. The present invention addresses these problems and deficiencies characteristic of the prior art and conventional apparatus used to supply aircraft bleed air.
Another aspect of this invention concerns the engine air driven starter. Air starters are conventionally air powered turbines mounted to the engine accessory gearbox. The starter turbine rotates at very high speed and drives the engine through a planetary gear system during engine acceleration to just below idle speed. Once the engine lights it begins to develop its own power and, at a speed below idle, accelerates away from the starter. An overrunning mechanical clutch allows the starter to disengage and then the starter air is shut off and the starter turbine comes to rest. Within a very narrow flight profile of the aircraft, the starter can sometimes be used for emergency engine relight, but only at conditions where the windmill speed of the engine is low enough that the starter clutch can be engaged without damage due to what is referred to as crash engagement.
Engine starters can not be used during normal aircraft cruise conditions; where the only means for relight is from the freely windmilling engine. One advantage of the present invention is that it permits operation of the air starter during all aircraft flight conditions thereby avoiding the delay in engine relight which can be associated with flight conditions unfavorable for fast windmill relights. Therefore engine designers are looking to solve the problems of in flight engine restart with respect to the relatively narrow available flight profile and crash engagement of the engine starter.
Mechanically powered means for reducing boundary layer drag of various aircraft parts such as wings, nacelles, and aircraft tail assemblies have been proposed in the past and in patent application Ser. No. 07/489,150 entitled "AIRCRAFT ENGINE STARTER INTEGRATED BOUNDARY BLEED SYSTEM", invented by Samuel Davison, filed Mar. 6, 1990 and assigned to the same assignee and in a patent application Ser. No. 07/531,718 entitled "GAS TURBINE ENGINE POWERED AIRCRAFT ENVIRONMENTAL CONTROL SYSTEM AND BOUNDARY LAYER BLEED", invented by the same inventor of this patent, filed on Jun. 1, 1990, and assigned to the same assignee, both incorporated herein by reference. A patent application Ser. No. 07/531,734 entitled "GAS TURBINE ENGINE FAN DUCT BASE PRESSURE DRAG REDUCTION", invented by the same inventor of this patent, filed on Jun. 1, 1990, and assigned to the same assignee, and incorporated herein by reference proposes a mechanically powered means of bleeding boundary layer air and reducing the drag of the aircraft by introducing at least a portion of the pressurized bleed air into the fan duct of the engine to reduce the base drag of the duct.